This disclosure relates to a gas turbine engine, and more particularly to a gas turbine engine component having a cooling hole that reduces or excludes a downstream diffusion angle.
Gas turbine engines typically include a compressor section, a combustor section and a turbine section. During operation, air is pressurized in the compressor section and is mixed with fuel and burned in the combustor section to generate hot combustion gases. The hot combustion gases are communicated through the turbine section, which extracts energy from the hot combustion gases to power the compressor section and other gas turbine engine loads.
The combustion gases generated during operation of the gas turbine engine are typically extremely hot, and therefore the components that extend into the core flow path of the gas turbine engine may be subjected to extremely high temperatures. Thus, air cooling arrangements may be provided for many of these components.
For example, airfoil and platform portions of blades and vanes may extend into the core flow path of a gas turbine engine. These portions may include cooling holes that are part of a cooling arrangement of the component. Cooling air is communicated into an internal cavity of the component and can be discharged through one or more of the cooling holes to provide a boundary layer of film cooling air at the outer skin of the component. The film cooling air provides a barrier that protects the underlying substrate of the component from the hot combustion gases that are communicated along the core flow path.